This invention relates to thermal barrier coatings for components exposed to high temperatures, such as the hostile thermal environment of a gas turbine engine. More particularly, this invention is directed to a composition and method for repairing a thermal barrier coating that has suffered localized spallation due to thermal fatigue and stress, poor coating processes, coating defects, localized damage, impact damage and other mechanical damage, or defective coating areas on new or used parts.
Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of nickel and cobalt-base superalloys. Nonetheless, when used to form components of the turbine, combustor and augmentor sections of a gas turbine engine, such alloys alone are often susceptible to damage by oxidation and hot corrosion attack and may not retain adequate mechanical properties. For this reason, these components are often protected by an environmental and/or thermal-insulating coating, the latter of which is termed a thermal barrier coating (TBC) system. Ceramic materials, and particularly yttria-stabilized zirconia (YSZ), are widely used as a thermal barrier coating (TBC), or topcoat, of TBC systems used on gas turbine engine components. These particular materials are widely employed because they can be readily deposited by plasma spray, flame spray and vapor deposition techniques. A commonly used type of TBC is a coating based on zirconia stabilized with yttria, for example about 93 wt. % zirconia stabilized with about 7 wt. % yttria This general type of TBC has been reported in such United States patents as U.S. Pat. Nos. 4,055,705, 4,328,285, and 5,236,745, which are incorporated herein by reference. Such TBC coatings have a relatively rough surface and do not provide adequate heat energy reflection in frequency ranges for certain applications. In addition, application of certain TBC coatings requires use of apparatus having a controlled atmosphere or vacuum. Accordingly, such coatings and methods cannot be effectively utilized in field repairs.
To be effective, TBC systems must have low thermal conductivity, strongly adhere to the component, and remain adherent throughout many heating and cooling cycles. The latter requirement is particularly demanding due to the different coefficients of thermal expansion between ceramic topcoat materials and the superalloy substrates they protect. To promote adhesion and extend the service life of a TBC system, a bond coat is often employed. Bond coats are typically in the form of overlay coatings such as MCrAIX (where M is iron, cobalt and/or nickel, and X is yttrium or another rare earth element), or diffusion alumninide coatings. During the deposition of the ceramic TBC and subsequent exposures to high temperatures, such as during engine operation, these bond coats form a tightly adherent alumina (Al2O3) layer or oxide scale that adheres the TBC to the bond coat.
The service life of a TBC system is typically limited by a spallation event brought on by thermal fatigue and stress, coating defects, mechanical damage, wear, and the like. Accordingly, a significant challenge of TBC systems has been to obtain a more adherent ceramic layer that is less susceptible to spalling when subjected to thermal cycling. Though significant advances have been made, there is the inevitable requirement to repair components-whose thermal barrier coatings have spalled. Though spallation typically occurs in localized regions or patches, the conventional repair method has been to completely remove the thermal barrier coating, restore or repair the bond layer surface as necessary, and then recoat the entire component. Prior art techniques for removing TBC""s include grit blasting or chemically stripping with an alkaline solution at high temperatures and pressures. However, grit blasting is a slow, labor-intensive process and erodes the surface beneath the coating. With repetitive use, the grit blasting process eventually destroys the component. The use of an alkaline solution to remove a thermal barrier coating is also less than ideal, since the process requires the use of an autoclave operating at high temperatures and pressures. Consequently, conventional repair methods are labor-intensive and expensive, and can be difficult to perform on components with complex geometries, such as airfoils and shrouds. As an alternative, U.S. Pat. No. 5,723,078 to Nagaraj et al. teaches selectively repairing a spalled region of a TBC by texturing the exposed surface of the bond coat, and then depositing a ceramic material on the textured surface by plasma spraying. While avoiding the necessity to strip the entire TBC from a component, the repair method taught by Nagaraj et al. still requires removal of the component from the engine assembly in order to deposit the ceramic material, and further requires the use of plasma spraying apparatus to effect the repair.
Moreover, existing sprayable TBC materials require some type of post drying or firing in order to be stabilized prior to high temperature use, and therefore are ineffective for in field, in situ repairs. Tape materials require an autoclave to apply, and are thus not feasible for in situ repairs. While plasma sprayed materials do not all require post-deposition heating, such materials have much rougher finishes, and cannot be applied in the field for in situ repairs without spraying powder throughout the rest of the engine (which requires a major cleaning step prior to subsequent engine operation).
In the case of aircraft turbine engines and large power generation turbines, removing the turbine from service for repairs results in significant costs in terms of labor and downtime. For these reasons, removing components having TBCs that have suffered only localized spallation is not economically desirable. As a result, components identified as having spalled TBC are often analyzed to determine whether the spallation has occurred in a high stress area, and a judgment is then made as to the risk of damage to the turbine due to the reduced thermal protection of the component that could lead to catastrophic failure of the component. If the decision is to continue operation, the spalled component must typically be scrapped at the end of operation because of the thermal damage inflicted while operating the component without complete TBC coverage. Additionally, some newer TBCs utilize a smoothing layer over the TBC for better heat rejection and air flow. Currently, there is no known way at present to replace this smoothing layer having a very smooth finish on damaged TBC.
Accordingly, it would be desirable if a repair method were available that could be performed on localized spalled areas of TBC on turbine hardware in field and in situ, without necessitating that the component be removed from the turbine, so that downtime and scrappage are minimized.
It would also be desirable to repair a smoothing layer on a damaged TBC in a manner that restores the very smooth finish of the smoothing layer, as well as restoring the heat rejection and airflow properties of the smoothing layer.
It would further be desirable to provide an improved smoothing coating for repair of damaged TBC that is easy to apply by brushing and spraying in situ.
The present invention provides a chemical composition and method for repairing a thermal barrier coating on a component that has suffered due to spallation, fatigue, stress damage, poor coating processes, mechanical damage, or wear of the thermal barrier coating. The use of the methods and compositions is particularly applicable for in-field repair of TBC coatings having spallation damage with a spall damage depth of between about 1 to about 7 mils (0.001 inch to 0.007 inch).
The composition demonstrates thixotropic properties as a result of its selected components, and particularly the inclusion of a nano-sized fumed ceramic material. The amount of fumed ceramic material such as alumina, titanium dioxide ,or silicone dioxide is provided so as to impart thixotropic properties to the coating composition. The thixotropic properties are very important in allowing the coating to be applied in a non-controlled environment such as an on-wing turbine engine assembly. Even though the coating dries relatively quickly depending upon solvent selection and solvent content, time for polymerization and other stabilizing chemical and physical interactions can be 8 hours or longer. The thixotropic nature of the coating composition allows it to be applied to a surface by a variety of processes without running, slumping or sagging while it dries, and further allows the coating to be worked as necessary. As used herein, the term thixotropic refers to a property of a material composition that enables it to flow when subjected to a mechanical force such as a shear stress or when agitated and return to a gel-like form when the mechanical force is removed. This definition is consistent with the definition of thixotropy as set forth in Hawley""s Condensed Chemical Dictionary (Thirteenth Edition) and the Encyclopedia Britannica. This property allows the coating to be applied in a production or field repair environment to surfaces having complex geometries, including but not limited to spalled areas of TBC coat ed components such as turbine blades and shrouds, without exposing the underlying component surface as a result of slumping, running or dripping of the coating. These important thixotropic properties allow the coating to be applied to a surface by any one of a number of processes such as spraying, dipping, brushing etc. The applied coating will not flow due to the effects of gravity such as by slumping, running, or dripping after application. However, the coating will flow if it is subjected to a mechanical shear stress, allowing it to be worked, if so desired. Thus, during the early stages of the drying period the coating can be worked if necessary. Of course, the ability to work the coating will be gradually diminished during the curing period, which is dependent on the curing of the binder, up until curing is complete. These properties allow the coating to overcome problems of dripping and running experienced with other coatings, which leave portions of the substrate uncoated.
Once applied and permitted to cure for about 8 hours at room temperature, the composition is stable and the resulting dry coating is ready for high temperature use. Thermal shock data on as-dried material was capable of withstanding temperatures of about 2000xc2x0 F., and IR reflection improvements were noted. Additionally, upon full curing, such as after heating by engine operation , the coating exhibits lower roughness or Ra values than TBCs applied by other methods.
As a result of the foregoing properties, the coating composition demonstrates adhesion to a wide variety of substrates, including but not limited to bare metal, grit blasted metal and coated metal, coated metals and coated ceramics of all types, TBCs, and many high temperature composites. The composition and method of the present invention allow on-wing, in-field, in-situ repair of TBC defects in smooth coat finish, and requires no post treatment heating or firing to stabilize before high temperature use.
The present invention provides for in-situ repair of a TBC coating without the need for post-coating heat treatment commonly required among known TBC repair methods. Indeed, the composition a l lows for use of room temperature or xe2x80x9ccoldxe2x80x9d spray application methods to effect a repair of the TBC coating. Alternatively, elevated temperatures can be utilized, such as heat lamp, heat blanket treatment, or heat gun treatment to accelerated dry and curing time without adversely affecting the desirable properties of the repair coating.
The method of applying the chemical composition of the present invention preferably involves spray application at room temperature, without the need to apply heat to cure the composition to effect the repair. In this embodiment, after cleaning the surface area of the component exposed by the localized spallation, the chemical composition is applied in a fine spray. The fine spray may be created by known aerosol and air spray means, including but not limited to aerosol assist canisters, air-pressurized canisters, high velocity, low pressure (xe2x80x9cHVLPxe2x80x9d) spray guns, airless sprayers, and other known air spray or liquid spray apparatus.
Regardless of the spray delivery system used, the spray propels the chemical composition against the affected area of the TBC to form a thin layer of ceramic and polymer composition. The solvent carrier then evaporates to leave a thin protective layer of binder and ceramic materials on the surface area of the component. The coating is permitted to dry at ambient temperature, preferably for at least about 8 hours. Drying and curing continues over time, and is further accelerated when the part is exposed to a heat source such as seen in a turbine engine flowpath. Drying may be accomplished at room temperature.
Optionally, the coating is heat treated soon after coating deposition to accelerate the dry and curing time. In any case, the coating is fully cured in situ by the engine heat which causes the binder to decompose or react to yield a ceramic-containing repair coating that covers the surface area of the component, and t hat comprises the ceramic powder in a matrix of a material formed when the binder is reacted at high temperature for a sufficient period of time. The binder is preferably a ceramic precursor material that can be converted immediately to a glass or ceramic by heating, or allowed to thermally decompose over time to form a glass-ceramic or ceramic repair coating.
According to the invention, each step of the repair method can be performed while the component remains installed at ambient temperature, e.g., in the flowpath assembly of an idle gas turbine engine. Within 8 hours after the final spraying step, at ambient temperature, the turbine engine can resume operation at which time the heat generated by operation of the engine will cure the applied coating to produce a ceramic. However, this time can be accelerated when drying and curing are optionally accelerated by application of heat treatment.
In view of the above, it can be appreciated that the invention overcomes several disadvantages of prior methods used to repair thermal barrier coatings. In particular, the method of this invention does not require the thermal barrier coating to be completely removed, nor does the invention require removal of the component in order to repair its thermal barrier coating. As a further advantage, the repair process does not require a high temperature treatment, since the repair coating exhibits sufficient strength to withstand engine operation, and is cured or fired by engine operation to form a ceramic/glass ceramic coating system.
Additionally, this coating provides a lower roughness finish with improved heat rejection due to lower energy transmission. As a result, minimal downtime is necessary to complete the repair and resume operation of the turbine engine. The invention can be used in any gas turbine component having a TBC, such as aircraft engines and turbines for electrical power generation. In the case of power generation turbines, the cost of completely halting power generation for an extended period in order to remove, repair and then reinstall a component that has suffered only localized spallation is avoided. Also avoided is the need to decide whether or not to continue operation of the turbine until the spalled component is no longer salvageable at the risk of damaging the component and the turbine.
Other objects and advantages of this invention will be better appreciated from the following detailed description. Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.